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At least 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Blade root torsional dampers to reduce hub loads

A new method of reducing helicopter rotor hub loads and marginally improving rotor performance by the introduction of large values of blade root torsional damping is presented. Basic theoretical considerations imply that these benefits in hub loads can come about by changes to the blade elastic torsional deflection. This basic theory was analytically verified by using a fully coupled aeroelastic rotorcraft analysis as applied to a modern, articulated rotor blade, namely that of the Sikorsky S-76. From an implementation standpoint, a rotor-based torsional damping device may be more practical than one that involves a major portion of the blade span. Also, a root-based device may allow for the retrofitting of existing helicopter rotor blade/hub configurations.

Kottapalli, Sesi↗
NASA NTRS1992

Fully integrated aerodynamic/dynamic optimization of helicopter rotor blades

This paper describes a fully integrated aerodynamic/dynamic optimization procedure for helicopter rotor blades. The procedure combines performance and dynamics analyses with a general purpose optimizer. The procedure minimizes a linear combination of power required (in hover, forward flight, and maneuver) and vibratory hub shear. The design variables include pretwist, taper initiation, taper ratio, root chord, blade stiffnesses, tuning masses, and tuning mass locations. Aerodynamic constraints consist of limits on power required in hover, forward flight and maneuver; airfoil section stall; drag divergence Mach number; minimum tip chord; and trim. Dynamic constraints are on frequencies, minimum autorotational inertia, and maximum blade weight. The procedure is demonstrated for two cases. In the first case the objective function involves power required (in hover, forward flight, and maneuver) and dynamics. The second case involves only hover power and dynamics. The designs from the integrated procedure are compared with designs from a sequential optimization approach in which the blade is first optimized for performance and then for dynamics. In both cases, the integrated approach is superior.

Walsh, Joanne L.↗
NASA NTRS1992

Investigation of advanced counterrotation blade configuration concepts for high speed turboprop systems. Task 4: Advanced fan section aerodynamic analysis

The purpose of this study is the development of a three-dimensional Euler/Navier-Stokes flow analysis for fan section/engine geometries containing multiple blade rows and multiple spanwise flow splitters. An existing procedure developed by Dr. J. J. Adamczyk and associates and the NASA Lewis Research Center was modified to accept multiple spanwise splitter geometries and simulate engine core conditions. The procedure was also modified to allow coarse parallelization of the solution algorithm. This document is a final report outlining the development and techniques used in the procedure. The numerical solution is based upon a finite volume technique with a four stage Runge-Kutta time marching procedure. Numerical dissipation is used to gain solution stability but is reduced in viscous dominated flow regions. Local time stepping and implicit residual smoothing are used to increase the rate of convergence. Multiple blade row solutions are based upon the average-passage system of equations. The numerical solutions are performed on an H-type grid system, with meshes being generated by the system (TIGG3D) developed earlier under this contract. The grid generation scheme meets the average-passage requirement of maintaining a common axisymmetric mesh for each blade row grid. The analysis was run on several geometry configurations ranging from one to five blade rows and from one to four radial flow splitters. Pure internal flow solutions were obtained as well as solutions with flow about the cowl/nacelle and various engine core flow conditions. The efficiency of the solution procedure was shown to be the same as the original analysis.

Crook, Andrew J.↗
NASA NTRS1992

Investigation of advanced counterrotation blade configuration concepts for high speed turboprop systems. Task 4: Advanced fan section aerodynamic analysis computer program user's manual

The computer program user's manual for the ADPACAPES (Advanced Ducted Propfan Analysis Code-Average Passage Engine Simulation) program is included. The objective of the computer program is development of a three-dimensional Euler/Navier-Stokes flow analysis for fan section/engine geometries containing multiple blade rows and multiple spanwise flow splitters. An existing procedure developed by Dr. J. J. Adamczyk and associates at the NASA Lewis Research Center was modified to accept multiple spanwise splitter geometries and simulate engine core conditions. The numerical solution is based upon a finite volume technique with a four stage Runge-Kutta time marching procedure. Multiple blade row solutions are based upon the average-passage system of equations. The numerical solutions are performed on an H-type grid system, with meshes meeting the requirement of maintaining a common axisymmetric mesh for each blade row grid. The analysis was run on several geometry configurations ranging from one to five blade rows and from one to four radial flow splitters. The efficiency of the solution procedure was shown to be the same as the original analysis.

Crook, Andrew J.↗
NASA NTRS1993

Study of the capacitance technique for measuring high-temperature blade tip clearance on ceramic rotors

Higher operating temperatures required for increased engine efficiency can be achieved by using ceramic materials for engine components. Ceramic turbine rotors are subject to the same limitations with regard to gas path efficiency as their superalloy predecessors. In this study, a modified frequency-modulation system is proposed for the measurement of blade tip clearance on ceramic rotors. It is expected to operate up to 1370 C (2500 F), the working temperature of present engines with ceramic turbine rotors. The design of the system addresses two special problems associated with nonmetallic blades: the capacitance is less than that of a metal blade and the effects of temperature may introduce uncertainty with regard to the blade tip material composition. To increase capacitance and stabilize the measurement, a small portion of the rotor is modified by the application of 5-micron-thick platinum films. The platinum surfaces on the probe electrodes and rotor that are exposed to the high-velocity gas stream are coated with an additional 10-micron-thick protective ceramic topcoat. A finite-element method is applied to calculate the capacitance as a function of clearance.

Barranger, John P.↗
NASA NTRS1992

Unsteady blade pressures on a propfan: Predicted and measured compressibility effects

The effect of compressibility on unsteady blade pressures is studied by solving the three-dimensional Euler equations. The operation of the eight-bladed SR7L propfan at a 4.75 deg angle of attack was considered. Euler solutions were obtained for three Mach numbers, 0.6, 0.7 and 0.8, and the predicted blade pressure waveforms were compared with flight data. The comparisons show that in general, the effect of Mach number on pressure waveforms are correctly predicted. The change in pressure waveforms are minimal when the Mach number is increased from 0.6 to 0.7. Increasing the Mach number from 0.7 to 0.8 produces significant changes in predicted pressure levels. The predicted amplitudes, however, differ from measurements at some transducer locations. At all the three Mach numbers, the measured (installed propfan) pressure waveforms show a relative phase lag compared to the computed (propfan along) waveforms due to installation effects. Measured waveforms in the blade tip region show nonlinear variations which are not captured by the present numerical procedure.

Nallasamy, M.↗
NASA NTRS1993

Evaluation of MARC for the analysis of rotating composite blades

The suitability of the MARC code for the analysis of rotating composite blades was evaluated using a four-task process. A nonlinear displacement analysis and subsequent eigenvalue analysis were performed on a rotating spring mass system to ensure that displacement-dependent centrifugal forces were accounted for in the eigenvalue analysis. Normal modes analyses were conducted on isotropic plates with various degrees of twist to evaluate MARC's ability to handle blade twist. Normal modes analyses were conducted on flat composite plates to validate the newly developed coupled COBSTRAN-MARC methodology. Finally, normal modes analyses were conducted on four composite propfan blades that were designed, analyzed, and fabricated at NASA Lewis Research Center. Results were compared with experimental data. The research documented herein presents MARC as a viable tool for the analysis of rotating composite blades.

Bartos, Karen F.↗
NASA NTRS1993

Structural tailoring of aircraft engine blade subject to ice impact constraints

Results are presented for the minimum weight design of SR2 unswept blade made of (titanium/graphite-epoxy/titanium) sub s fiber composite. The blade which is rotating at high RPM is subject to ice impact. The root chord length, blade thicknesses at five stations, and graphite-epoxy ply orientation are chosen as design variables. Design constraints are placed on the behavior variables: local leading edge strain and root damage parameter (combined stress failure criteria) as a function due to ice impact, maximum spanwise centrifugal stress at the root of the deformed blade due to local damage, first three natural frequencies, and resonance margin after impact. The method of feasible directions is employed to solve the inequality constrained minimization problem. The effect of ice speed and the ice impact location on the final design are discussed.

Reddy, E. S.↗
NASA NTRS1992

Aerodynamic design of turbomachinery blading in three-dimensional flow - An application to radial inflow turbines

A computational method, based on a theory for turbomachinery blading design in three-dimensional inviscid flow, is applied to a parametric design study of a radial inflow turbine wheel. As the method requires the specification of swirl distribution, a technique for its smooth generation within the blade region is proposed. Excellent agreements have been obtained between the computed results from this design method and those from direct Euler computations, demonstrating the correspondence and consistency between the two. The computed results indicate the sensitivity of the pressure distribution to a lean in the stacking axis and a minor alteration in the hub/shroud profiles. Analysis based on Navier-Stokes solver shows no breakdown of flow within the designed blade passage and agreement with that from design calculation; thus the flow in the designed turbine rotor closely approximates that of an inviscid one. These calculations illustrate the use of a design method coupled to an analysis tool for establishing guidelines and criteria for designing turbomachinery blading.

Yang, Y. L.↗
NASA NTRS1991

Euler solutions to nonlinear acoustics of non-lifting rotor blades

For the first time a computational fluid dynamics (CFD) method is used to calculate directly the high-speed impulsive (HSI) noise of a non-lifting hovering rotor blade out to a distance of over three rotor radii. In order to accurately propagate the acoustic wave in a stable and efficient manner, an implicit upwind-biased Euler method is solved on a grid with points clustered along the line of propagation. A detailed validation of the code is performed for a rectangular rotor blade at tip Mach numbers ranging from 0.88 to 0.92. The agreement with experiment is excellent at both the sonic cylinder and at 2.18 rotor radii. The agreement at 3.09 rotor radii is still very good, showing improvements over the results from the best previous method. Grid sensitivity studies indicate that with special attention to the location of the boundaries a grid with approximately 60,000 points is adequate. This results in a computational time of approximately 40 minutes on a Cray-XMP. The practicality of the method to calculate HSI noise is demonstrated by expanding the scope of the investigation to examine the rectangular blade as well as a highly swept and tapered blade over a tip Mach number range of 0.80 to 0.95. Comparisons with experimental data are excellent and the advantages of planform modifications are clearly evident. New insight is gained into the mechanisms of nonlinear propagation and the minimum distance at which a valid comparison of different rotors can be made: approximately two rotor radii from the center of rotation.

Baeder, J. D.↗
NASA NTRS1991

Correlation of airloads on a two-bladed helicopter rotor

Airloads measured on a two-bladed helicopter rotor in flight, from the Tip Aerodynamic and Acoustic Test, are compared with calculations from a comprehensive helicopter analysis (CAMRAD/JA), and the pressures compared with calculations from a full-potential rotor code (FPR). The flight test results cover an advance ratio range from 0.19 to 0.38. The lowest speed case is characterized by the presence of significant blade-vortex interactions. Good correlation of peak-to-peak vortex-induced loads and the corresponding pressures is obtained. The results of the correlation for this two-bladed rotor are substantially similar to the results for three- and four-bladed rotors, concerning the tip vortex core size for best correlation, calculation of the peak-to-peak loads on the retreating side, and calculation of vortex-induced loads on inboard radial stations.

Hernandez, Francisco↗
NASA NTRS1930

Test of an adjustable pitch model propeller at four blade settings

This note describes tests of an adjustable blade metal model propeller, both in a free wind stream and in combination with a model fuselage, at four settings of the blades. The model propeller is designed for a uniform nominal pitch/diameter ratio of .7 and the blade settings used correspond to nominal pitch/diameter ratios of .5, .7, .9, and 1.1 at the .6 radius. The tests show that propellers of this type may be considerably changed in setting from the designed pitch angles and yet give excellent performance. The efficiency realized and power absorbed when blades are set at other than the designed angle, are little different than would be obtained from a propeller with uniform pitch equal to the mean pitch of the propeller under test.

Lesley, E P↗
NASA NTRS1939

Observations in Flight of the Region of Stalled Flow over the Blades of an Autogiro Rotor

The flow over the inner halves of the rotor blades on a Kellet YG-1B autogiro was investigated in flight by making camera records of the motion of silk streamers attached to the upper surfaces of the blades. These records were analyzed to determine the boundaries of the region within which the flow over the blade sections was stalled for various tip-speed ratios. For the sake of comparison, corresponding theoretical boundaries were obtained. Both the size of the stalled area and its rate of growth with increasing tip-speed ratio were found to be larger than the theory predicted, although experiment agreed with theory with regard to shape and general location of the stalled area. The stalled region may be an important factor in both the rotor lift-drag ratio and the blade flapping motion at the higher tip-speed ratios. The method of study used in this paper should be useful in further studies of the problem, including the reduction of the size of the region.

Bailey, F J , Jr↗
NASA NTRS1939

Propeller Tests to Determine the Effect of Number of Blades at Two Typical Solidities

Propellers with equal total blade area, but with different numbers, were tested at Stanford University. The tests show generally that, for equal total blade area, propellers with the larger number of blades absorb the greater power and, provided hubs have equal drag, develop the higher efficiency. It is shown that the differences found are in agreement, qualitatively, with what might be predicted from simple blade-element theory.

Lesley, E P↗
NASA NTRS1951

Determination of blade-to-coolant heat-transfer coefficients on a forced-convection, water-cooled, single-stage turbine

Blade-to-coolant convective heat-transfer coefficients were obtained on a forced-convection water-cooled single-stage turbine over a large laminar flow range and over a portion of the transition range between laminar and turbulent flow. The convective coefficients were correlated by the general relation for forced-convection heat transfer with laminar flow. Natural-convection heat transfer was negligible for this turbine over the Grashof number range investigated. Comparison of turbine data with stationary tube data for the laminar flow of heated liquids showed good agreement. Calculated average midspan blade temperatures using theoretical gas-to-blade coefficients and blade-to-coolant coefficients from stationary-tube data resulted in close agreement with experimental data.

COOLING - GAS TURBINE SYSTEMS↗
NASA NTRS1951

Some effects of blade trailing-edge thickness on performance of a single-stage axial-flow compressor

A set of modified NACA 65-series blower blades designed for axial inlet velocity, high inlet Mach number, and high blade loading was investigated for trailing-edge thickness of 0.015, 0.030, and 0.045 inch to determine the effect of trailing-edge thickness on single-stage axial-flow-compressor performance. Trailing-edge thickness effects were small except at the highest tip speed investigated (915 ft/sec). Trailing-edge thickness up to 30 percent of maximum blade thickness were used without sacrifice of performance of NACA 65-series blades.

AIRFOILS - THICKNESS↗
NASA NTRS1951

Determination and Use of the Local Recovery Factor for Calculating the Effectiveness Gas Temperature for Turbine Blades / Jack B. Esgar and Alfred L. Lea

In an in experimental investigation of local recovery factors for a blade having a pressure distribution similar to that of a typical reaction-type turbine blade, it a was found that the recovery factors were essentially independent of Mach number, Reynolds number, pressure gradient, and position on the blade surface except for regions where the boundary layer was probably in the transition range from laminar to turbulent. The recommended value of local subsonic recovery factor for use in calculating the effective gas temperature for gas turbine blades was 0.89.

HEAT TRANSFER, CASCADES↗
NASA NTRS1951

Blade-to-coolant heat-transfer results and operating data from a natural-convection water-cooled single-stage turbine

Blade-to-coolant heat-transfer data and operating data were obtained with a natural-convection water-cooled turbine over range of turbine speeds and inlet-gas temperatures. The convective coefficients were correlated by the general relation for natural-convection heat transfer. The turbine data were displaced from a theoretical equation for natural convection heat transfer in the turbulent region and from natural-convection data obtained with vertical cylinders and plates; possible disruption of natural convection circulation within the blade coolant passages was thus indicated. Comparison of non dimensional temperature-ratio parameters for the blade leading edge, midchord, and trailing edge indicated that the blade cooling effectiveness is greatest at the midchord and least at the trailing edge.

COOLING - GAS TURBINE SYSTEMS↗